text:
What would be the challenges to taking off and landing a typical passenger jet at FL300?
Suppose it was somehow possible build a runway at FL300. Yes, doing so would be incredibly expensive, and it would be a mammoth civil engineering task, and the downsides would probably by far outweigh its advantages in practice; but for the time being, let's say that the desire to do it exists, and the issue of money has been resolved (as if by magic).
Let's also ignore the issue of getting the passengers, cargo, etc., to and from the actual aircraft, and focus on the aircraft operation itself.
Having constructed this runway-in-the-sky, what difficulties, compared to a similar runway built at ground level, would pilots and airplanes face when using it to "land" (if we can even use that term here) and take off, respectively, with a typical passenger jet (turbofan) aircraft?

It has been suggested that this is a duplicate of Would building an airport at 30,000 ft save money? (which also, in the question body, allows for one at 10,000 ft). I believe that this question is not a duplicate of that one because that question focuses on the cost of building such an airport and the savings in fuel expenditures because airplanes would already be near cruising altitude, while this question focuses on the challenges of landing at and taking off from an extremely-high-altitude airport while specifically ignoring the economic aspects.

Possible duplicate of Would building an airport at 30,000 ft save money?

This was well covered by comments and answers on another recent question.

@abelenky I don't think I agree. The difficulties that the pilots and airplanes would face were mentioned in answers to that question, but really only in passing, I think.

Maybe some S.H.I.E.L.D. agent is around and can tell how their Helicarrier (a flying aircraft carrier in the Marvel Cinematic Universum) works. (Sorry, couldn't resist.)

@abelenky Adding to TannerSwett's comment above, the linked question seems to me to be about monetary savings from such an extremely high-altitude airport, while this question is about actually operating an airplane into and out of such an airport if one could be built (but excludes "ground" operations such as baggage handling, passengers, etc., as well as the monetary aspect). Related, yes; duplicate, I don't really see how. Looking at Bianfable's answer to this and their answer to the other one (where it's highest voted and accepted at this time), the two answers are completely different.

Why 30,000 feet?  Why not ask about landing at 40, or even 50 thousand feet?

These are the main challenges I can think of:

High takeoff and landing speed: To generate enough lift for takeoff in the thinner air at FL300 (gravity has almost not decreased, so required lift stays almost the same), very high speeds are required. Combine this with less thrust from the engines available and you need very long runways. It should still be possible to accelerate and takeoff eventually since FL300 is still below the maximum altitude most jet airliners can climb to (coffin corner).
Tires: The tires used on aircraft have a maximum speed which would be exceeded when taking off or landing. Other tires could be developed, but it is a major challenge to make these work in the cold temperatures (rubber does not like freezing temperatures, that is why Mars rovers do not use rubber tires).
Cold Temperature: It is cold at FL300, around 220K on average or around -50°C. This is not only problematic for the tires, but also for fuel, which would be constantly close to its freezing point. A heat exchanger between e.g. engine oil and fuel would be mandatory to operate an aircraft from here.
Wind: Wind speeds can easily reach 150kt at FL300. The runway would have to be perfectly aligned with the wind to have a pure headwind component. Otherwise, the crosswind would quickly exceed the limits for the aircraft.
Engine Failure: FL300 is above the maximum altitude for OEI (One Engine Inoperative) flight, which implies one cannot return to this airport after an engine failure. An aircraft would be forced to drift down to a normal airport for landing.

Not to mention that at -50C an engine that's shut down for an hour or so during unloading/loading/refueling would see some issues with lubrication.  Engine heating blankets, fuel tank heaters, hot tankers for fueling, etc.

@ZeissIkon Good point, pretty much everything would have to be heated by external sources while the engines are off.

Can you even achieve positive climb at takeoff and go-around configuration at FL300?

@Jimmy you can climb in clean configuration during normal operation, so why not? You just need typical cruise speeds on the ground.

Cruise configuration is probably the only configuration here that allows for positive rate. So we are talking about Fast & Furious 6 length of runway here. Should probably also mention that Dom's crew would all require oxygen bottles at this altitude.

Re "High takeoff […] speed": Out of curiosity: what would be the minimum speed to maintain level-flight at FL300 for a mid-sized aircraft (e.g. A320)? At sealevel I figured it is around 150kt (±20).

@DJClayworth I agree, but the OP said "ignore the issue of getting the passengers, cargo, etc., to and from the actual aircraft", so I ignored everything related boarding and baggage handling.

@PerlDuck I am not sure what exactly the minimum speed would be, but the 737 FCOM says best holding speed at FL300 is 178KIAS for 40t total weight and 263KIAS for 85t total weight. The minimum is probably not far below. Note that these are indicated speeds, the TAS would be in the range from 250 to 400 knots at these altitudes.

"so I ignored everything related boarding and baggage handling" Indeed, the whole point of that restriction in the question was to attempt to narrow the question down sufficiently to be (a) answerable and (b) actually about aviation.

Assuming that this airport really is floating in the sky or built on the top of a very large tower, wouldn't it be enough for takeoff to just push the plane over the edge with just enough horizontal momentum for it to have just enough lift to not exceed its Vne while diving down, gaining speed, thus gaining lift, etc.? Still doesn't solve the landing problem though.

If you can point the runway straight into a 150kt headwind it alleviates the first two problems noticeably :)

Flight range would be limited
Aircraft would not be able to take off with full fuel, because of the difficulty of taking off in the high altitude.
In fact some planes could not take off at all at max weight, not even with unlimited runway, because with a full fuel load they are at "coffin corner" at that altitude, and incapable of (safely) attaining the altitude. Planes typically take off and fly 27,000, 29,000, 33,000, 35,000 etc. as they burn off fuel and become capable of the higher altitude.
Taking off, a too-heavy airplane would have to run off the end of the runway. Then nose down to reduce the severity of the immediate stall, and sacrifice altitude to gain speed and get to a more practical altitude.  The runway would need to end at a ramp arching down a cliff, which will make RTOs a lot more dicey.


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Why might a too-wide runway be a problem?
In the September 2020 issue of EAA's Sport Aviation, Budd Davisson writes

Runway width becomes a problem for some taildraggers when it gets to be less than 35-40 feet because of the lack of visibility.

Sure.  When you can't see the centerline, it's nice to have some lateral margin of error.
But later in the article (p. 74) in the caption of an overhead photo of a small airplane on a runway seven times wider than its wingspan, he writes

Oddly enough, super-wide runways can cause as many problems as very narrow ones, especially for taildraggers.

What might such problems be?  (Not wandering into the grass.  Even if you can neither look over the nose nor locate the runway's edges, if during takeoff or rollout you drift to one side, wouldn't you still notice the center stripe or the edge or the lights in time to correct?)
Do any NTSB reports mention a too-wide runway as the cause of an incident?

There's an old joke about the pilot and copilot landing and having to bring the plane to a very abrupt stop. "That's the shortest runway I've ever seen!" said the pilot. The co-pilot looked from side to side and said "yeah, it's also the widest ..."

The way I heard it, @Jason, was much less "politically correct", but it's still funny. :)

"Tower to Speedbird, you were a bit to the left of the centerline on that landing."  "Speedbird to Tower, yes, and my co-pilot was a bit to the right."

At the former Kelly Air Force Base, a taxiway had "TAXIWAY" painted in HUGH letters easily seen on final to the adjacent runway. Over the years too many had, or attempted to, land on the taxiway.

The issue has always been human perception.  Pilots are tasked with trying to make great landings.  One way they do that is with peripheral vision.  If the runway is too wide, they lose that extra clue on when to roundout and flare.  They may flare too early and stall the aircraft too high above the runway.  This perceptual clue is more important for tailwheel pilots than tricycle gear pilots.
From the FAA AIM 8-1-5.

Illusions Leading to Landing Errors.

(a) Various surface features and atmospheric conditions encountered in landing can create illusions of incorrect height above and distance from the runway threshold. Landing errors from these illusions can be prevented by anticipating them during approaches, aerial visual inspection of unfamiliar airports before landing, using electronic glide slope or VASI systems when available, and maintaining optimum proficiency in landing procedures.

(b)Runway width illusion. A narrower‐than‐usual runway can create the illusion that the aircraft is at a higher altitude than it actually is. The pilot who does not recognize this illusion will fly a lower approach, with the risk of striking objects along the approach path or landing short. A wider‐than‐usual runway can have the opposite effect, with the risk of leveling out high and landing hard or overshooting the runway.

Though it perhaps be mentioned that this is to some extent a matter of habit, e.g. judging flare by the perceived width.  If you occasionally land on dry lakes and such (or wet ones, I understand, though I have no personal experience of seaplanes), you learn to deal with the illusions.

The base I currently fly out of has a 10,000x150ft runway.  One of our emergency divert fields is 5,000x75ft.  The perception would be that it is the same as the runway we land on everyday, when in reality it is half the length.
In pilot training, we had a 13,000x300ft runway next to an 8,000x150ft.  I'd always find myself flaring high when landing on the 300ft wide runway.

@Keegan: Perhaps that's a consequence of not landing at enough different runways?  So you've unconsciously trained yourself to use the perception of runway width as a cue to flare.  So it's not that the runway width is inherently a problem, it's the habit that's been formed.

@jamesqf: In the context of safety, the human factor cannot be understated. Consider the opposite effect, of a pilot used to a wide runway landing at a narrower one. Can you see the CFIT in the making?

@dotancohen: Sure, and that's why I suggested becoming used to different sorts of runways so you DON'T get into the habit of using the perceived size as a cue.

When landing an airplane you get used to visual reference points to evaluate your position relative to the ground. Optical illusion are therefore quite common.

A narrower-than-usual runway can create an illusion that the aircraft is higher than it actually is, leading to a lower approach. A wider-than-usual runway can create an illusion that the aircraft is lower than it actually is, leading to a higher approach and greater slope.

On a very wide runway you might flare too early and even stall the airplane a few meters above ground. Taildraggers are quite tricky at landing in those condition  as you don't want to overshoot the 3 points pitch attitude of the aircraft or you will hit the tail first which is not designed to handle it.
The opposite might happen on narrow runway for nose wheel plane where you might miss your flare and touchdown with the nose gear first who might collapse under you. This risk adds obviously to the risk of hitting ground obstacle while flying a low approach.

Thanks for explaining why it's a problem "especially for taildraggers" as Budd wrote.

It's a problem for everyone, but consequences might be a bit worse in a taildragger ^^

At the former NAS Dallas, about 1/4 of the runway juts out into a lake on a peninsula just wide enough for a taxiway on one side.  My first time on short final I felt slightly unnerved with a "what's wrong with this picture?" I got used to it, but it recurred a couple of times.   The OLS meatball was my friend! Years later I was not surprised when Asiana 214 hit the  runway approach end at SFO for want of referencing the visual approach lights. Who hasn't landed with one VASI light red? They were not "just a little low."

As a pilot, switching from my training on a 50 ft runway to landing frequently on a 150 ft wide runway was slightly challenging.
The FAA describes two illusions:
Wide Runway
Looks closer on final and you will tend to float and flare high.
Narrow Runway
Looks farther away on final and you will tend to approach at a higher rate of descent.

Is it feasible to land away from the centerline, and closer to one side of the runway?  Separately, is that even permitted ?

@Criggie Surely it ispossible and authorised assuming your not entierly missing the runway, but the real question is why would you do it ? Flying in close formation is the only valid reason I can find, and we often do it with aicraft landing on each side of the runway to reduce the time and space between landings. Othervise, the centerlinegive a perfect reference, give you the more security in case of crosswind landings where the plane light woble left to right or any other problem.

Does the size of runway markings help with this? Probably not an urgent problem, but in a theoretical world maybe a fixed-size marking further out on the runway, so that it can be a reference point for height.

Larger runways are made for larger airplanes - therefore, the markings are more complex and designed for a longer TDZ(touch down zone) for aircraft moving a lot faster than 65 knots.

The issue with the FAA illusions is that as a trainee pilot you learn a mental "picture" of what your whole visual field should look like, and how it should be located relative to the edges of your windshield, when flying a visual approach. If the runway is a different width from what you expect, your brain will misjudge the width and/or the length and/or your angle of descent, to make what you see match up with your internal "picture" of how you should be moving in three dimensions.

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What is the takeoff decision speed (V1) for a Gulfstream G500?
Can someone tell me what is the decision speed (V1) of the Gulfstream G500? Is it 200kt as shown in this picture?

Your decision speed also called V1 is the speed at which ou have to commit to taking off whatever happens because you won't be able to stop before the end of the runway after that. This speed is related to the airfield runway length, aicraft TOW, temperature and altitude and will change for each take-off. Therefore there is not a fixed value for a given aicraft.
For the G500 I did find this review

Today our take-off weight with three persons on board and 11500 pounds  of jet fuel (one-third of the max fuel) is 60 000 pounds and our computed speeds are V1: 118 kt for a rotation speed (Vr: 123) and a take off safety speed V2: 129 kt.

And this other one

The G500 was relatively light, loaded with 12,000 pounds of fuel for our flight, which was scheduled for 1.5 hours. Takeoff weight was 59,863 pounds, well below the 79,600 mtow. The avionics calculated V1 at 116, Vr 118, V2 132, and Vref in case of immediate return at 136 knots.

As you can see the value changes slightly but V1 seems to stay around 120kts. Your 200kt speed limit and the altitude and range limit related seems more likely to be linked to the 91.117(b) FAA regulation

Unless otherwise authorized or required by ATC, no person may operate an aircraft at or below 2,500 feet above the surface within 4 nautical miles of the primary airport of a Class C or Class D airspace area at an indicated airspeed of more than 200 knots (230 mph.). This paragraph (b) does not apply to any operations within a Class B airspace area. Such operations shall comply with paragraph (a) of this section.

The 200kt limit appears to be to comply with 91.117(b): there's a speed limit of 200kts below 2500' within 4nm of a class C or D airport. Those numbers match the departure numbers in the OP's picture.

You might be absolutely correct, is it FAA regulation?

Yes, it's FAA. It's in 14 CFR 91.117.

Okay, flying in Europe I didn't knew, I'll edit my answer to stick to actual facts, thanks

There is no one speed for V1. It’s dependent upon the loadout of the jet, as is Vr, Vref, etc.  To be honest for that kind of a small business jet, a V1 of 200 KIAS is way too fast; the value is probably in the neighborhood of 110-120 KIAS with Vr around 130-135 KIAS.

200 KIAS for V1 is too fast for pretty much any aircraft. Even Concorde had typical V1 speeds of well below 200 KIAS (although VR was usually in that range).

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Writing Research: What stick sensation would you expect in an Su-15 with a total hydraulic flight system failure?
The title pretty much says it all. I have a very specific research question in relation to a writing project I'm working on.
The main character is in an Su-15 performing low flying military training, and has a total hydraulic flight system failure (which I understand is extremely rare) and crashes into the side of a mountain. I envision moments before, the pilot pulling back on the stick as hard as he can, with no authority, or give in the stick, resulting in the plane smashing into the mountains ahead. However, I'm not entirely sure this would have been the case and need to be as accurate as possible for this project.
Thanks

The pilot would most likely eject before crashing into the side of a mountain.

You need to specify an actual aircraft @JohnP.Good. Some soviet fighters of the era (Mig-15, Mig-17) didn't have hydraulic controls to fail in the first place. Those with hydraulic controls will vary significantly in characteristics after a hydraulic failure.

@GdD, thanks for the suggestion. The aircraft is an Su-15.

The SU-15 was an interceptor, not a fighter, it's not realistic it would be flying low level.

Regardless of the role of the jet, low level flight would be trained.

Su-15 (as well as all Sukhoi interceptors since Su-7) had a completely irreversible control system with no manual override. This means, 100% of torque was produced by hydraulic boosters, and the force on the control stick was simulated with a special variable spring loading mechanism.
So, a theoretical total hydraulic failure would result in loss of control, while the stick would move 'normally' - just without any effect.
Su-15 had 4 hydraulic systems, with 2 of them used for flight control. The most anticipated reason for dual failure would be loss of the engines; for that, there was an electric pump in one of the systems. In case of control problems, the pilot would check and ensure that the emergency pump was on. This pump provided only a limited amount of control (and for a limited time).
Now I'll have to challenge the premise a bit. First, Su-15 was an interceptor, and low-altitude flight (even training) would be quite unusual for it (though not impossible). Its lowest altitude of intercept was 2 km, and the main emphasis was on intercepting higher altitude targets. Its radar was unable to track targets on the earth surface background. It was not meant for terrain following, so even if it was flying at a low altitude, it would be fairly clear of immediate obstacles. Which gives the pilot at least a few seconds to react.
Second, the pilot would, of course, eject. Even though in emergency situations (esp. unanticipated emergencies) people do all sorts of things, pilots, and esp. pilots of aircraft with irreversible control, know that loss of control = eject. It's pointless to 'fight' with the stick. However, as I said, in all likelihood there would be a few seconds to think. In most situations (like the engines failure) you won't have an immediate loss of control: the failure is fairly gradual. This would see the pilot trying to ascertain what was happening. Difficulties with control at low altitudes often require reducing the airspeed.1 But a catastrophic sudden loss of control is a clear signal to eject.

1 By the way, early Su-15 had pitch control problems at transonic speeds at low altitude due to insufficient booster torque. It felt like no response (or poor response) at higher G manoeuvres. This things could be resolved by reducing speed.

In an aeroplane with hydraulically powered, irreversible flight controls, the controls are very hard to move with no hydraulic pressure. The controls have a bit of movement around the servo valve travel, a couple of centimeters at the stick. The unpowered hydraulic actuator acts as a damper, and the pilot must exert quite a bit of force to deflect the stick past the servo valve travel.
On ground, the elevator/ailerons will deflect when in the servo valve stops, at great effort. In the air, the aerodynamic hinge moments resist stick deflection, and the pilot may pull what they want, without this effort resulting in any surface deflection.
In flight, the effects of gradual pressure loss are not much noticeable at first. Flight control deflections (and associated aerodynamic hinge moments) in normal flight are relatively small, and the remaining pressure in the last hydraulic system will be able to cope with these until the underpowerd actuator runs out of authority. Which also happens when the pilot attempts a drastic pull-up manoeuvre when near the ground.
Total hydraulic pressure loss in an irreversible hydraulic aeroplane is a catastrophic event, and there will always be redundancy built into the aeroplane. Multiple hydraulic systems, each with multiple pumps, with urgent warning signals tothe pilot when one of the system parameters is out of the normal range. So the pilot won’t be taken by complete surprise, but aeroplane response can suddenly run out when all pressure is lost. A suspenseful scenario.


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What is the difference between the various automated airport weather observing systems?
Automated airport weather observations can be produced by several types of equipment: AWOS-A, AWOS-1, AWOS-2, AWOS-3, and ASOS are the ones I've seen mentioned. What are the differences between these reporting systems? Are the older systems still prevalent, or have they been mostly converted to the new systems?

ASOS is basically a replacement for AWOS with more reporting parameters, and the first units were installed in 1991.  According to the Automated airport weather station Wikipedia page, each different type has the following parameters:
AWOS

AWOS A: barometric pressure and altimeter setting (in inches of    Mercury).
AWOS I: wind speed and wind gusts (in knots), wind    direction (from which the wind is blowing) and variable wind    direction (in
  degrees of the compass), temperature and dew point (in    degrees
  Celsius), altimeter setting and density altitude.
AWOS II:    all AWOS I parameters, plus visibility and variable visibility (in    miles).
AWOS III: all AWOS II parameters, plus sky condition (in    oktas), cloud ceiling height (in feet), and liquid precipitation
  accumulation (in inches).
AWOS III P: all AWOS III parameters, plus    precipitation type (rain, snow and sometimes drizzle) identification.
AWOS III T: all AWOS III parameters, plus thunderstorm detection (via    a cloud-to-ground lightning detector).
AWOS III P/T: all AWOS III    parameters, plus precipitation type identification and thunderstorm    detection.
AWOS IV Z: all AWOS III P/T parameters, plus freezing rain    detection via a freezing rain sensor (Note: this configuration used
  to be called AWOS III PTZ).
AWOS IV R: all AWOS III P/T parameters,    plus runway surface condition.
AWOS IV Z/R: all AWOS III P/T    parameters, plus freezing rain detection and runway surface    condition.

ASOS

They generally report all the parameters of the AWOS-III, while also
  having the additional capabilities of reporting temperature and dew
  point in degrees Fahrenheit, present weather, icing, lightning, sea
  level pressure and precipitation accumulation.

What are the practical reasons a pilot would need to remember these differences, or know what type a particular airport has?  Just tune to the weather frequency prior to approach and take note of whatever details that are spoken...

@bovine: Don't ask me, I didn't ask the question.  :-)

@bovine while I agree that you'd just tune in prior to approach or take-off, it might be helpful to know what information is not given for each one.  Most of it the difference have to do with relaying the type of precipitation observed, so theoretically, you could be surprised by snow or freezing rain if you thought that the report should include it (but didn't mention it, so you thought you'd be fine) but rather the station was unable to report it due to lack of sensors.

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How do oceanic lateral offset procedures work?
I was reading a forum for professional pilots and ran across a thread on lateral offset procedures during oceanic flight (apparently having the FMS offset the aircraft from the centerline of an oceanic track). How does lateral offset work? Is it required for all oceanic aircraft or optional for some operators? Is it explicitly assigned by ATC or is it an implicit part of oceanic flight? What is its purpose, and why isn't it used in non-oceanic airspace?

On high-altitude North Atlantic flights, planes can fly zero, one, or two miles to the right of the centerline of your assigned track. The offset is randomly selected. This is called Strategic Lateral Offset Procedures.
It's used in the oceanic flights or flights over remote land where there is typically no radar service, and ATC cannot "see" traffic separation. It provides, on average, additional horizontal separation between planes reducing the likelihood of midair collision and high altitude wake turbulence (which can occur as a result of RVSM).
Strategic lateral offsets can be authorized in en-route oceanic or remote continental airspace. The decision to apply a strategic lateral offset is the responsibility of the flight crew, and only where offsets have been authorized by the ATS and when the aircraft is equipped with automatic offset tracking capability.
https://www.faa.gov/about/office_org/headquarters_offices/ato/service_units/enroute/oceanic/documents/WATRS_Plus/DOC_4444_16.5_SLOP_25Feb2011.docx

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When are winds given with respect to true vs. magnetic north?
Wind information can be reported by various sources (ATIS, METAR, TAF, spoken on the radio, etc). I was taught that officially some of these sources are relative to magnetic north and others to true north (although practically this isn't always followed). Which sources are supposed to use which reference system?

The general rule is:
If you read it, it's true. If you hear it, it's magnetic.
All charts and textual sources (METAR, TAF, winds aloft, surface analysis charts, etc) use true north as the reference.
ATIS/AWOS/ASOS broadcasts, or any information a controller gives you over the radio, is magnetic.

Wind direction broadcast over FAA radios is in reference to magnetic north.

AIM Section 7-1-11 (page 7-1-26 in the 5/26/16 edition)
One exception to the "if you hear it" rule is that a FSS briefer will read you the winds referenced to true north, since they're just reading you the charts/textual information.
(This is at least true in the United States... other countries may vary in some instances)

It's the same in Europe. Wind information communicated by the tower when a landing/takeoff clearance is given is in degrees magnetic. METAR/TAFs are in degrees true. Not that it matters much because magnetic variation is almost zero in Western Europe

I believe one very important difference is TWEBs, or Transcribed Weather Broadcasts.  Because they are an automated broadcast of a written report, they are reported in true headings, just like a written report, even though you technically hear it.

Oops, I down-voted before I researched my answer completely.  It appears we were both wrong.  The AIM seems to indicate, as I expected, that the ASOS/AWOS data is the same as METAR, as I had expected.  What I was wrong about them both being magnetic.

@2NinerRomeo you might want to research your answer... a little more completely ;)

Any idea what the reason is behind this? Seems like it would be easier/safer if everything was in one or the other. Thx!

This answer provides some insight: https://aviation.stackexchange.com/questions/18918/why-are-magnetic-headings-used-for-designating-runways

Sorry for the necro but I believe that the D-ATIS (digital ATIS) is also an exception to this rule. It also reports wind direction relative to magnetic north. It is the textual version of the ATIS that is shown in the cockpit and basically serves the same function as a spoken ATIS so that makes sense.

In the United States, the principal guide to weather products is AC 00-45G "Aviation Weather Services".  As of this writing it appears Change 1 is current, published July 2010.  Citations given here are to that revision.  The Airman's information manual has information on AWOS and ASOS.
Here are some examples:

METAR: True 3.1.3.5
AWOS & ASOS: True AIM 7-1-11 Figure 7-1-5
TAF: True 7.2.2.5
WA (Winds Aloft): True 7.4.2
UA (Pilot Reports): Magnetic 3.2.1.9 (I'd suppose you get whatever the pilot reports.)
Surface Analysis Chart: True 5.1.2.3.5

When calling flight service, the briefer will have available the same information products that you can reach on aviationweather.gov.  I expect they will know the direction reference for the information they are giving you.

When you say AWOS/ASOS to a pilot, they think of the automated radio report, where wind is given as magnetic, not true. I think you're getting mixed up because the ASOS equipment is also used to create METARs and other textual products, where wind is reported as true. Quote from http://www.nws.noaa.gov/asos/pdfs/aum-toc.pdf (section 3.2): "Wind direction is reported relative to true north in
the METAR/SPECI message.... Wind direction is reported
relative to magnetic north in the computer-generated
voice messages"

Also, AC 00-45G is cancelled. AC 00-45H is current.

AIM Figure 7-1-5 is about decoding a textual METAR generated from ASOS/AWOS. It has nothing to do with the verbal radio broadcasts. See the note on page 7-1-26: "Note- Wind direction broadcast over FAA radios is in reference to magnetic north." (taken from the 5/26/16 AIM).

The way I remember it is that if it relates to an aerodrome its magnetic, because runway numbers are in magnetic and the pilot needs to know how the wind relates to the runway in use so the amount of headwind / crosswind can be assessed.
If it relates to en-route then its in true, because flight planning is done in true (with conversions to the correct magnetic heading to set on the compass at each step).

This is what I thought until I looked it up.  I was gravely disappointed because this is how I thought it should be.

Another good general rule
Everything the weather man tells you it's true ... the rest is magnetic

This seems opposite to the previous answer.  Are you talking about the TV weather man?  That doesn't seem useful for aviators.

@PJNoes How did you get TV weather from this?  It's an attempt to memorize that weather information is provided in True North, while operational (atis, tower) is Magnetic. "True" in this case meaning True North and not Accurate, Correct,etc.. hence the funny pun... well, attempt of, anyway...

I grew up when the only weather was from the weather man on TV.  I got the funny pun, but the previous answer is easier for me to remember.  No offence meant.

If the weatherman wrote it, its true.  If he spoke it or tells it, its magnetic.  Written is True.

when is the weatherman speaking to you using magnetic north ? ... is what I'd like to understand ....


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What is the difference between planned true airspeed and actual true airspeed?
My electronic E6B does calculations for both planned and actual true airspeed (yielding different results). I was told the difference has to do with assumptions about temperature, but the manual doesn't give any details. What is the difference between the two?

When you do the calculation for actual, are you using CAS or IAS to do the true calculation? That's a step that's often overlooked (or impossible)!

From the manual for an electronic E6B which seems to be similar to yours:

PLANNED TRUE AIRSPEED
This function is used to calculate true airspeed for preflight
  planning. It will compute the true airspeed in knots and Mach  number
  and density altitude, given the pressure altitude,  temperature, and
  calibrated airspeed in knots.

and

ACTUAL TRUE AIRSPEED
This function calculates true airspeed, Mach number and  density
  altitude given pressure altitude, indicated temperature  in Celsius
  and calibrated airspeed.

The manual doesn't give any more information on the difference between the two calculations so we can only guess...
Both methods use pressure altitude and calibrated airspeed. One clue is that the "planned" version uses "temperature", while the "actual" version uses "indicated temperature".
So my guess is:

you use the "actual" version in flight, and it expects you to input Total Air Temperature (TAT), read from the gauge in the airplane, which it then converts to Outside Air Temperature (OAT) to do the density altitude and true airspeed calculation.
You use the "planned" version for flight planning, where the temperature you have available is the OAT.

The TAT is the OAT plus the rise in temperature due to the air being brought to rest relative to the airplane.
Since the TAT is a bit higher than the OAT, the "actual" version should give slightly lower density altitude and a slightly lower true airspeed.


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What happens if you change power setting with constant mixture setting?
For a given Cruise Setting: Manifold set / RPM set / Mixture set according to P.O.H. values: what happens to the Mixture if you were to reduce power or increase Power by varying throttle? Does a change in power setting always require a change in mixture to maintain desired stoechiometric ratio (e.g.: best power)? Why? I somehow know it is the case but cannot figure out why. Thanks

All things equal, the throttle setting has no meaningful impact on the stoichiometry of the fuel-air mixture, it just changes the volume of that mixture that's allowed to flow into the cylinders.
There are reasons to enrich the mixture beyond stoichiometric efficiency that have to do with assisting the engine with managing its temperature, but this isn't about power, it's about not breaking the engine - allowing uncombusted fuel to exit the engine, carrying heat with it (fuel has more density, and is therefore a useful thermal mass).
Under cruise conditions, however, if you increase throttle, you're going to gain altitude (unless you re-trim for a higher speed).  The reverse if you cut throttle.  Fuel-Air-Mixture is, as you observe, stoichiometric but higher altitude air is lower pressure and therefore lower density - requiring fewer molecules of fuel to achieve stoichiometric equilibrium.
So if you reduce power, you're going to need to enrich the mixture sooner or later, because you're now descending, not because you reduced power per se.

So the bottom line is that “No” a mixture adjustment is not required when power is changed, it it is often desirable.


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What is considered to be a flight into known Icing?
I am aware that flight into known icing is prohibited if your aircraft is not suitable to fly in icing conditions (unless e.g.: FIKI,...). Given the following example I am wondering if it would be legal to takeoff in Luxembourg in these conditions if your cruising level is well above the cloud tops: Cloud Layer with Cloud tops +/- FL050 and w/ Freezing Level FL040+ and light icing to be shown on a icing forecast / report. What indications are best to use to predict icing? Would this report indicate flight into known icing (I guess Yes, but I am not sure) thus making a takeoff "illegal"?, although you would only briefly pass the narrow layer of icing? The safest answer is obviously to not takeoff(although I do believe it would present no hazard in this situation), but I am simply questioning this from a legal point of view? Thanks!

I see this is tagged easa-regulations, @tTed are you also interested in FAA based answers (thats what mins article covers)?

For FAA regulations, see this question.

This is a wonderful question! The regulations are vague around it. According to your situation presented above, it is certainly illegal. AIM 7-1-22 defines Known Icing Conditions as "Atmospheric conditions in which the formation of ice is observed or detected in flight."
I urge you to check out this letter which is a response to AOPA from the FAA regarding known icing conditions:
https://www.govinfo.gov/content/pkg/FR-2007-04-03/pdf/07-1620.pdf
Here is a quote from that letter: "The NTSB has held that known icing conditions exist when a pilot knows or reasonably should know of weather reports in which icing conditions are reported or forecast."
The FAA argues that if a pilot willingly flies into known icing conditions, they could take legal action as per 14 CFR §91.13 (no reckless or careless behavior).

Welcome to aviation.SE! We already have a question that covers this from the FAA point of view.

Just talked with some experienced pilots and they all concluded that this is a case into know icing as you expect icing to occur at temperatures below 0 degrees celsius thus climbing through a layer (visible moisture) known to be above freezing level  is definitely "flight into known icing", especially from a legal point of view, no matter if you only pass it during a few seconds.

Keep in mind that passing through that 1000ft layer of icing conditions takes waaaay longer than just a few seconds, unless you are flying a fighter jet. For a general aviation acft a cruise climb would be something in the 500 to 1000 ft/minute range, so should the conditions really be bad, you'd have plenty of time to accumulate a nice crust on your wings... Temperature below 0 degrees celsius does not necessarily mean conditions are icing. You need sufficient humidity for that too.


text:
Why does propeller slip (efficiency) change during the flight (fixed-pitch)?
I was wondering how the efficiency of a fixed-pitch propeller changes during different stages of flight and why it does so. I suppose when we are talking about propeller efficiency we are actually talking about propeller slip. So my follow up question is what causes propeller slip to
change, what conditions make a propeller lose efficiency. I am aware that it has to do with Pitch, but I can't figure out why efficiency changes throughout the flight. THANKS

Are you looking for a detailed engineering explanation, or just a big - picture understanding?

Big Pitcure understanding

The reason that efficiency is reduced during phases of flight that are flown at high airspeeds is because the angle of attack between the propeller and relative wind is reduced. If an aircraft is stationary on the ground, its propeller is taking a larger "bite" of air because there is no relative wind. When there is more air moving past the prop, it takes a smaller "bite" of air, therefore producing less thrust and being less efficient. The advantage of a constant speed prop is that you can change the pitch of the prop depending on your airspeed, so it maintains a constant angle of attack, and is more efficient in every phase of flight. I'm not totally sure that's what you are asking, but that's what I got :)

Pitch is essentially how much the propeller would move forward in a complete rotation if it was by itself. You can now rotate the propeller at any rate (read RPM) you can achieve. So if you pitch is 10 cm, and your RPM is 1000, then in 1 minute it would 10000 cm in 1 minute ,ie 100 metres, giving a velocity of 1.66 m/s. If the propeller could move at that speed there would be zero slip, but it in reality it is attached to a shaft and has certain "loading". While moving through air, it experiences drag as well as generates thrust, and since it has a body attached to it, that also experiences drag and cumulatively a significant drag. So the actual distance travelled is may be 60 m, giving a speed of 1 m/s. (40m is the slip here). The drag is a closely related to velocity (for entire aircraft) and rpm of propeller (for the prop). Now the propeller pitch being fixed, and the velocity changing through out the flight, the velocity would be changing in the entire flight regime and consequently drag and therefore slip.
Hence you can imagine, there would be velocity for which the slip is the least. This would be the design speed, where efficiency is maximum (the peak). To either side of this speed efficiency would be lower.


text:
What would be the challenges for a pilot taking off into a headwind of just under the clean stall speed?
If a strong and fairly steady headwind was coming down the runway at just under the stall speed for a clean configuration/weight, what challenges would a pilot face? Or would it be fairly easy? Is this something that happens often?
Edit: This is the video that made me ask.

Just to confirm, the windspeed is just under the stall speed of the aircraft?  Because for a light civil GA plane this would be a gale force storm.  For an airliner or military jet, it would be hurricane force winds.  Pilots don't fly in winds like that.  It isn't taking off in a headwind that's the problem, it's taxiing without flipping over that makes it dangerous.

Here's an example of what happens in a strong wind near stall

@MichaelHall It was pretty late at night when I asked this lol.

Robert has answered the question quite well, but I will elaborate:
Taking off into a strong headwind is not difficult as such, but as Robert (and John as well, in less subtle way :) pointed out, taxiing in strong wind, however steady it might be, is a real pain it the seat of the pants. Since the question is about taking off, lets assume the plane would (magically) get to the end of the runway safely. The takeoff itself would then be something along the lines of what can be seen for example in this video:
Youtube: STOL Competition - World Record Shortest Landing 9 Feet 5 Inches
(Yes the title says "landing", but there are plenty of takeoffs to be seen)
Nothing special really, except you are airborne in a blink of an eye. What might be troublesome to an inexperienced pilot, is that the visual cues are weird, so airspeed would have to be monitored with great care as referencing the surroundings would be useless. Also: great wind speed at ground level guarantees there will be turbulence, so maintaining appropriate airspeed is critical, and lifting off the plane should be "decisive" to avoid being pushed back on the ground by shifts in wind. The question states a wind speed just under clean configuration stall speed, so you would not use flaps, since as soon as you would lower them, you would lift off.
As for the part of "Is this something that happens often?": No. Outside the realm of STOL contests, during normal operations: this does not happen. Regulations and POH limitations set aside, you just could not get on the runway in one piece.
P.S. It is worth noting that majority of STOL contest planes are taildraggers, so this gear configuration certainly is not a disadvantage when taking off or langing at windspeeds closing stall speed.

"Fairly steady" wind is the only wording that would make it even remotely considerable, with risks of being ground looped or blown over while taxiing.
One scenario may be a light GA plane taking off from a moving aircraft carrier.
Tricycle gear would be a tremendous advantage, as this would help keep the wing at a very low angle of attack to the wind.  This allows the plane to remain on the ground during its takeoff run long enough to establish a safety margin in airspeed before the plane "rotates" to an angle of attack sufficient to generate enough lift to fly.
One must be aware, although this is doable, gusts make it extremely dangerous.  Far better to wait for better weather.
Even with no wind, the safe rotation airspeed will be higher than the clean stall speed.

A taildragger would be just as able to keep the AoA down by simply raising the tailwheel off the ground. Perhaps even more so than a tricycle geared plane, prop clearance permitting.

Yeah but forget get about trying to taxi in any direction other than directly into wind.

Oh absolutely

The main challenge would be the long walk back to the airport, assuming you could still walk, after your airplane was blown like a tumbleweed into the next county.  In other words, it's simply not something you would even try.  Only very experienced or foolish pilots will even go out in winds approaching 50% of stall.


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How can I modify a fuselage to increase the payload?
I'm designing a regional turboprop for a university assignment. The design requirements are very similar to the Dash 8 Q400 in terms of passenger capacity (90 pax), speed, range, runway length requirements. The difference is the payload mass required is 35% greater (8,500kg to 11,500kg).
So I've sized my wings and my engines and control surfaces and flaps and they're all bigger. My issue now is the fuselage. Since the passenger capacity is the same I have less motivation to increase the dimensions.
Increasing the diameter for the sake of more internal volume seems inefficient; I don't need more aisle and armrest space. On the other hand, increasing the length (without increasing diameter) would mean a heavier fuselage structure.
Even once the fuselage is lengthened, where do I put the cargo? For the Q400, cargo holds are right at the aft and just before the cockpit. But if this is my solution I'll be shifting the centre of gravity and creating even greater bending moments.
My ideas so far are to create a  near c.g. cargo hold that would likely split seating or to utilise space available under the floor of the cabin. I haven't seen any designs that support these ideas though (for narrow bodies).
I tried to look for twin-engine turboprops with similar payloads, the only one I found was the C-27J Spartan (military transport), but it's practically designed for cargo.
Could someone give me some wisdom on how I should proceed? I've done quite a bit of reading and research but can't resolve this aspect of the design.
Here's the 74 seat cabin layout for the Q400.

Do you want to increase the payload volume, or the payload mass?  You will get different answers for each.

If your number of passengers is given, so is their weight. ICAO has standard passenger masses which increase with trip range. What does the additional payload consist of? As @jamesqf says, it makes a difference if it's styrofoam blocks or bricks.

@Jamesqf, Payload mass to 11,500kg (from Q400's 8,500kg).

@PeterKämpf Thanks I hadn't seen the standard ICAO masses before. The assignment brief specifies 90 passengers @ 106kg (!), but the aircraft must be able to carry a payload of 11,500 kg total. Strangely the design driver is for short-haul flights of 600km, I don't know why the passengers need so much luggage.

Thank you for the details. Should you assume ICAO's 161 kg/m³ payload density, or is another value given? Splitting the cabin in two is not recommended (emergency exits!), so make the fuselage wider and gain more space under the cabin floor.

@Peter Kämpf: Plus you have to consider how dense the cargo is going to be.  If it's dense, you'll want it somewhere near the center of gravity.  If it's light but bulky, you have a quite a bit more freedom.

You really only have two options; stretch or swell.
The Q400 already has a "combi" configuration option with 50 seats forward and a cargo hold taking up the aft third of the cabin, so if you want to just add another couple of plugs, one (smaller plug) forward of the wings and one aft, to get 74 seats plus a large cargo area, you can do that and just adjust the structural requirements to deal with the even longer flying pencil that results.  You can have a cargo hold aft and pax cabin forward with the appropriate weight and balance requirements and limitations for loading pax vs cargo same as any other split combi layout.
In the case of the Q400 itself, its configuration is already at the extreme limit of what you can do with fuselage stretches on that sort of 4-across tube diameter, having started with 36 seats in the original.  The plane suffers from tail clearance problems because of this, due to limits on how much the gear legs could be extended to account for the long tail (landing with less than full flaps - there are a couple of intermediate landing flap settings - has a high risk of tail strikes and pitch attitude must be very carefully managed in the landing flare).
I think you'd get a better aircraft by going with "swell" and increasing the tube diameter enough to get a 5 across cabin (around 2 feet), with 2 on the left side and 3 on the right. A 74 pax 5-across cabin would only need 15 rows, leaving the rest free for cargo, and this extra width plus shortened passenger area might even allow you to shorten the fuselage a bit, leaving some stretch-ability potential in your initial design (you might even get bonus points for taking this into account).

Thank you, very informative. I guess that's part of the reason why people say the Q400 is difficult to land. I didn't realise that tail strikes put a limit on the coupling length (?) of the aircraft but it's definitely something I'll bring up in my report. I'm likely going to use the second configuration you recommended but move the seating backwards a little bit so luggage can be distributed both at the aft and the front. I posted a picture of the Q400 cabin layout in my post now, something like that just with 5 abreast seating.

That would work if you have the forward cargo hold on one side only.One issue with large aft cargo areas is you have to allow for a loaded cargo hold with an empty cabin on the ground, and you may be forced to move the gear location farther aft, requiring a larger horizontal tail, or worse, include a tail strut.So consider that in C of G calcs.Tail engine jets have this problem with no pax aboard cuz of the wt of the engines, and place the gear farther aft than necessary for the normal loaded C of G requirement, to avoid a tail strut, and requiring a larger elevator for rotation on TO.


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How is the Dash 8 Q400 so fast?
The Q400 has a cruise speed of 360kts, and on flightradar24 most Q400s that I looked at reached similar speeds during cruise. The ATR 72 has similar specifications yet only flies at 280kts. I'm currently designing a regional turboprop as part of a university assignment and I can't find a drag polar that gives 360kts as anywhere near the optimum speed for L/D ratio, 244kts was the cruise speed for my initial configuration. Is the Q400 likely flying at a much higher, less efficient speed? Why would this be done? Or has better technology allowed it to fly efficiently at this speed?
Edit: According to a pilot on this flight sim forum they used to cruise at around 200kts. This video shows Flybe (budget airliner) cruising at 240kts, FL240. Maybe the flights I've seen on flight radar are operating under different parameters at the moment, with fewer passengers and lower oil prices maybe 360kts is the best speed.

Are you sure the drag polar is at cruise altitude? 244kts IAS at FL300 may very well be 360kts TAS.

@EfeBallı It doesn't really change your point, but service ceiling for the Q400 ist FL270. Most cruise below FL250 because they don't have oxygen masks (for passengers).

@Efe Ballı thought you might be right, 360kts TAS is about 241kts EAS at FL250. I'm fairly sure it's 244TAS though. Here are my optimal climb and cruise TAS https://i.imgur.com/dDP1nmc.png C_do = 0.0273 and C_di = 0.0283. This is a plane designed for a heavier payload than the dash8.

The designers of Q400 made the plane light enough to be fast. The major weight savings come from main ldg gear, which is so light it collapses approximately on every other landing. Sorry... I'll get  mi coat...

@Jpe61 hahaha, I'll be sure to include that in my design report.

Oh dear god don't, you'll fail the course

I'm sure I read somewhere that the Q400 cruises at 400kts TAS, it's a major selling point of the Q400 as it cruises almost as fast as a jetliner. I don't have any data as proof, but I don't believe 400 kts is too far off its best range speed.
@Jpe61, not sure about weight but it surely saves a fair bit of drag to retract the LG into the nacelle and eliminate fuselage protrusions for LG housing. 100% agree on LG fragility

"ATR 72 has similar specifications".  That is not a true statement and negates your entire question. The Q400 is simply a faster aircraft.

@EfeBallı yes, the Dash 8 series is way more streamlined than the ATRs. I'm pretty these planes were designed with rather different goals in mind, so the speed of Dash 8 is not just about burning more juice.

Here are the cruise performance tables for the Q400 at the heaviest and lightest weight (click to view larger):
Source: from the Dash-8 Q400 performance manual, which came included with a simulator product I purchased. They probably got their data from a QRH (Quick Reference Handbook), which usually includes such tables in the In-Flight Performance section.
At ISA conditions at FL250, the long range cruise is flown between 209 and 188 KIAS (depending on weight) and the high speed cruise is flown between 230 and 243 KIAS. The corresponding TAS at FL250 (assuming ISA conditions) are:

188 KIAS: 281 KTAS
209 KIAS: 312 KTAS
230 KIAS: 343 KTAS
243 KIAS: 363 KTAS

That last value (for high speed cruise at light weight) seems to agree quite well with the max. cruise speed of 360 KTAS, which one usually find for the Q400. But best range speed at that weight would be around 188 KIAS (281 KTAS), quite a bit slower. At higher weight, the speed differences between best range and high speed cruise are smaller though.
At typical weights the Dash-8 will still be faster than the ATR 72-600. I don't have exact performance tables for that aircraft, but Wikipedia says cruise speed is 280 KTAS. So yes, the Dash-8 is built for higher speed. The biggest difference is the choice of engine: PW150A with 3782 kW for the Dash, but PW127M with only 1846 kW for the ATR, more than a factor two difference! While this difference probably causes a higher best range speed for the Dash compared to the ATR, in absolute fuel usage the ATR is still more economical.

Thanks. Do you mind sharing where you got these tables?

@RoryMcDonald I always include a link to the source if it is publicly available. In this case, I couldn't find a public source. I got the tables from the Dash-8 Q400 performance manual, which came included with a simulator product I purchased. They probably got their data from a QRH (Quick Reference Handbook), which usually includes such tables in the In-Flight Performance section. Maybe you can find a QRH on Google somewhere...

I embedded the source info into the answer, in case a comment clean up ever accidentally sweeps that away.


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Why are vortices produced at wingtips a source of drag, while vortices from the leading edge (harnessed by delta wings) good?
From my understanding, wingtip vortices produce a downward force on the aircraft, effectively increasing drag. Meanwhile, when considering leading-edge vortices on delta wings, we only consider the reduced pressure created by the vortices to be a good thing.
Why is the negative pressure not considered for the wingtips? Why isn't the downwash considered for the leading-edge vortices?

https://aviation.stackexchange.com/questions/21069/what-is-vortex-lift has a great answer (and a reference to why non-delta wings are different in the comments): "One main condition for vortex lift generation is high sweepback angles."

You ask the right question. Given the amount of misinformation out there, I can understand why you are confused.
First the tip vortices: They are neither a source of downward force not do they cause drag. They are only a consequence of lift creation by wings of limited span. This answer tries to explain this in more detail. If, however, you mean the wake vortex, then yes, you can view them as a source of drag. But still not of a downforce.
Now to the vortex of a delta wing at high angle of attack. This is caused by flow separation at the leading edge and produces a low pressure area over the wing. While the tip vortex picks up its rotation speed while air is flowing over the wingtip, and then continues into the free airstream of the wake, the vortex of the delta wing directly affects pressure on the upper side of the wing. Without it, the wing would stall at a much lower angle of attack, so it helps a delta configuration airplane to land at a much lower speed. But it also causes a lot of drag: Flow separation at the leading edge means that leading edge thrust is lost and the lift force is essentially orthogonal to the wing surface. Given the high angle of attack, the drag of a delta wing with vortex lift is very high (lift force times the tangent of the angle of attack).
But this is not all bad: It helps to burn off energy quickly which is useful during approach and landing. Certainly, you do not want to rely on vortex lift during cruise: Then the lift of a delta wing will be created much like it is on unswept wings, tip vortices included.

Interesting. So would this video be incorrect? To save you some time, just skip to 1:03. It says that the inner part of the tip vortex is deflecting air downwards, but because you said " They are neither a source of downward force not do they cause drag", that would kind of disprove that video. Thanks for your time.

@Wyatt As I have explained countless times, the video has it backwards. The vortex is the consequence of downwash. Please, please, don't believe random videos but try to use your own brain.

I see. It is a little confusing, as the explanation in the video of the inner tip vortex causing downward flow feels believable. (Downward moving air moving the air beside it downwards makes sense) - Perhaps it’s just a matter of definition that I’m confused about?

@Wyatt While flowing over the wing, air gets that downward push which produces lift. When leaving the wing, it continues moving down and the vortex forms. What is so confusing? The idiots who make those videos!

I think I was just set on believing the tip vortex caused the downwash, but I can see where it would actually be a result. Thanks again.

Sorry that I am asking another question, but I just found this page from NASA (a source I can trust), in which it says - "The vortices produce a downwash of air behind the wing which is very strong near the wing tips." It says this about halfway through the first paragraph. This made me confused, as I now thought the tip vortices didn't cause any downwash, but were the result. Maybe this is saying something different that what I was originally asking?

@Wyatt Better trust nobody, not even me. That NASA page was probably written by some intern who copied something found on the internet. Please start to think for yourself and don't blindly believe stuff you find on the interwebs!

Oh okay haha. I thought NASA was a source to always believe, but I guess even their articles have false information!


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Would English Electric Canberra benefit from a larger aspect ratio?
Considering high altitude performance while carrying a given payload, wouldn't Canberra be able to climb higher with a wing of same area but higher aspect ratio? Why was Canberra built with a relatively large chord?

From Wikipedia:

The Canberra had its formal origins in a 1944 requirement issued by
the Air Ministry for a successor to the de Havilland Mosquito. This
requirement, the initial revision being E.3/45, sought a
high-altitude, high-speed bomber, which was to be equipped with no
defensive armament.

The relationship with the Mosquito followed in its development:

During the early stages of design, the aircraft had grown from being
roughly the same size as the Mosquito to being around double its
weight.[6][clarification needed] Although jet-powered, the Canberra
design philosophy was very much in the Mosquito mould, providing room
for a substantial bomb load, fitting two of the most powerful engines
available, and wrapping it in the most compact and aerodynamic package
possible, an example being a leading edge formed of a single sheet of
light alloy wrapped around to 40% of chord, sitting on Redux-bonded
stiffeners through which the ribs were passed, the panels secured with
adjustable eye-bolts, enabling a highly accurate wing profile to be
maintained from the leading edge to main spar without any external
joints or fastenings.[11] Also in line with the Mosquito philosophy,
the Canberra by design dispensed with defensive armament, which had
historically proven unequal to fighter aircraft, and the resulting
performance gain permitted the Canberra to avoid air-to-air combat
entirely.

As to its high altitude capabilities, a Canberra B.2 set a world altitude record, flying at 63,668 ft (19,406 m) on May 4, 1953.
Being a bomber, the added efficiency of a larger aspect ratio wing was not seen as a necessity. While it may have added to its range, the longer wings would add structural weight that would likely reduce bomb loads.
But to your question, Martin was building the aircraft in the US under license where it was designated the B-57 Canberra, including several RB-57s used for reconnaissance. That eventually led to the RB-57D:

The RB-57D was built strictly as a high-altitude reconnaissance
aircraft. It originated in a December 1952 USAF study funded by the
Wright Air Development Center for a turbojet-powered special
reconnaissance aircraft with a radius of 2,000 nautical miles (3,700
km; 2,300 mi) that could operate at altitudes of 65,000 feet (20,000
m). Subsonic performance was considered to be acceptable and it was
felt that no defensive armament would be needed. The RB-57D was unique
and set the stage for high-altitude reconnaissance operations in the
rarefied air of the stratosphere. Preliminary specifications were
prepared by the Air Force on 27 March 1953. The project was carried
out in high secrecy. It was known as Weapon System MX-2147, and the
code name was Bald Eagle.
The basic B-57 Canberra design was used as a starting point for the D
model, but there were several significant changes incorporated. The
most obvious was the greatly enlarged wing. The wingspan was
lengthened by 42 feet (13 m) to 106 feet (32 m) overall. The chord of
the wing was also increased, which combined with the increased length
gave the wing a very high lift capability.

New higher performance J57 engines and new lightweight wing construction to keep the weight down helped provide the enhanced performance. While they performed for a number of years, eventually, there was a need to upgrade resulting in the RB-57F:

The RB-57F was the result of an Air Force Big Safari requirement for a
high-altitude reconnaissance platform with better performance than the
existing and similar RB-57D, some of which had been grounded as a
result of wing spar failures.
The prototype RB-57Fs incorporated many major changes from the RB-57D,
the most obvious of which was an enlarged computer-designed wing to
enable it to operate at extreme altitudes. The wing had a span of more
than 122 feet (37 m), which was 16 feet (4.9 m) longer than the RB-57D
and nearly twice the length of the B-57B fuselage on which it was
installed. Extensive use was made of aluminum honeycomb wing panels in
the wings that bonded outer and inner aluminum skins to a honeycomb of
aluminum and fiberglass.[N 3] All control surfaces had tightly sealed
gaps in order to reduce drag, and there were no wing flaps. In
addition, the size of the empennage was redesigned so that the
vertical stabilizer had nearly twice the area of that of the standard
B-57B. Its height was increased to 19 feet (5.8 m) and the width
increased, improving longitudinal and asymmetric control for greater
stability at very high altitudes (up to 80,000 feet (24,000 m)).

The RB-57F also included an upgrade to P&W TF-33-P-11 turbofans, which had more than double the thrust of the original J65 engines. Four aircraft were configured as WB-57F high altitude weather aircraft. Three remain in service with NASA.
Everything in aircraft design is a compromise. The original Canberra did what it was required to do. It evolved from the Mosquito with the engines that were available at the time.
Longer and higher aspect wings were added later to support new missions, but required new wing manufacturing techniques and higher power engines. And that process was repeated again. So, yes, it could benefit from the higher aspect ratio wing and did - eventually.


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Does tailless delta configuration retain any advantage over conventional configuration with straight wings at high subsonic speeds?
For a model aircraft designed for cruise at say 0.75 Mach, do delta wings offer any benefit over straight wings. Is drag saving from elimination of the tail large enough to consider a delta configuration?

Does "straight" mean "non-swept"? Question would be improved by editing to clarify this-- as far as I can tell, all answers posted so far are actually comparing delta wings to swept wings.  Suggested edit: in body of question replace "straight wings" with "straight (non-swept) wings".  I'm not 100% sure that this is the actual intent of questioner so will leave it up to him/her--

Delta wings were studied in the early jet age for use in passenger planes but lost out to the conventional configuration. Avro in Great Britain studied a civilian version of their Avro Vulcan bomber. It was called type 722 Atlantic and could not compete with conventional designs. The tail drag of a conventional design is small when compared to the additional friction drag of the much larger wing area of a delta.
A tailless delta will need a much larger wing area for the same landing speed and have higher induced drag in cruise due to the shorter wingspan. This is slightly offset by a much better structural efficiency of the delta wing, so the empty weight of a delta airplane is quite a bit lower than that of comparable conventional designs. If ease of construction is important, the delta is far superior to a conventional design because it will not have the complex high lift devices which help the conventional design to get away with much less wing area.
Deltas shine in other areas: Their low lift curve slope makes them much better for low level flight because gusts result in smaller load factors. This is ideal for low level penetration of air defenses, but that is hardly something a passenger jet is used for. If you want to fly your model at low level and Mach 0.85, go with the delta wing. In dense air and at high speed, induced drag is minuscule, so that disadvantage of the delta falls away.
If you want to include flying wings in the comparison: A tailless straight wing at Mach 0.75 is a poor choice. There is little to compensate for Mach tuck in an overspeed event and the reflex airfoil is pretty much the opposite of what works well in transsonic flow, which is supercritical airfoils. The delta wins hands down. Its structural efficiency allows the use of thin airfoils and vortex lift gives it agreeable landing speeds even without the use of flaps.

The question refers to a conventional configuration with straight wing

@Abdullah yes, I'll edit the question

@Abdullah ... and I edit my answer.

The term "model" is one item preventing a resounding NO, but other wise evidence is strongly against the proposal.
The British V-bomber program of the 1950s produced 3 designs for high subsonic, long range (cruising) applications, the Vickers Valiant, the Handley Page Victor, and the delta Avro Vulcan.  The US had its similarly sized Boeing 707.
In terms of range, the delta Vulcan fell far short of the others.  Amazingly, the best design for cruising was not the 707, but the Handley Page Victor!  Politics, poor planning, lack of funds, and poor execution resulted in this advanced subsonic design being relegated to military use.
During operation Black Buck, it was the Victors who refueled the Vulcans on their way to the Falklands and back.  The greater strength of the Vulcan delta and its supersonic capability kept it in service longer, but it was the "ugly duckling" Victor design, with its cranked wing and wing root/fuselage jet engine air intakes, that could have given the 707 a run for its money.
The obsession of modern designers to save drag by eliminating tail volume makes about as much sense as a car saving weight by eliminating the steering wheel.
Trusting that a high subsonic "model" would have a Reynolds number greater than 10$^6$, the Handley Page or 707 design may be a good place to start.  For a slightly lower Mach specification, such as 0.55, a (relatively) tiny straight wing, such that seen on the DeHavilland Canada Dash 8 turboprop, will give best results for cruising, with traditional slats and flaps for landing.


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Why might a canard aircraft suddenly pitch down from a level flight?
I have a canard aircraft which has pitch stability issues that I'm trying to troubleshoot.
The aircraft has a characteristic tendency to depart level flight in sudden dive/ or generate a massive nose-down pitch rate that inverts the aircraft after which the aircraft tumbles and recovery is not possible. It seems to be set off when the pilot tries to give nose-down commands. I cannot figure out if this is purely a CG placement issue or if something is wrong with my wing/canard geometry or incidence.
The canard has an area of 20% of the wing, the moment arm is 2.4 times the MAC of wing and horizontal tail volume is -0.467.
EDIT 1: aircraft geometry

All dimensions in mm.
Canard airfoil: SD7062 (4.5 deg incidence)
Wing airfoil: E1230 (0 deg incidence)
Control surface on canard runs throughout the span. Control surface chord is 35% of canard chord.
Canard is mounted higher than wing by 165mm.

I hope this is a simulation you are talking about. If so, could you show some plots?

Immediate first guess is that your canard is stalling.  However that does not match with the sign of the input.  Is this a flying model, a computer simulation, or what?  If it is a naive simulation, the developer may not have anticipated a canard and there may be an error in the physics model.

@U_flow Sadly its not a simulation, Its a 25 Kg aircraft for my masters project.

what happens when you command pitch-up?

@nielsnielsen Shows expected behaviour, Nothing abnormal.

How does it look like? Which airfoil have you used? Where is the CG? Do you have a picture? How do control surfaces look like? I suppose the canard geometry has to be tuned a bit, maybe a different airfoil or aspect ratio will be needed

The canard may be interfering with the aft wing enough to affect trim.  When pitch down command is given, AoA on canard reduces, and all of a sudden the aft wing has much more lift (from reduction of interference)(I think Langley had the same problem).  A picture of your design would be very helpful.

@Mridul, do you have a data recording device onboard?

I would try increasing the incidence of the canard surface in increments and trying out each increment. When you apply nose down elevator, the effective incidence of the surface is reduced as the chord line is angled downward by the trailing edge moving up. Possibly the ND down elevator is reducing canard AOA to zero or even negative at some amount of TE up displacement, which would explain the tuck under, which is the canard/wing seeking a trim state where both surfaces are lifting in the same direction.

@Mridul aside from checking control throws up and down, you may try mounting the canard higher or lower relative to the main wing to rule out interference effects.  What type of airfoil are you using for main wing and canard?

@RobertDiGiovanni  Updated the aircraft geometry, I'm not at liberty to post real pictures of aircraft as yet

@RobertDiGiovanni Can you give me a reference for langley's aircraft?

@JohnK That is what I think is happening as well. It seems to happen close to 4-5 deg on up deflection. I can't figure out once the tuck under happens why do my control surfaces stop responding? Also can't figure out a possible remedy for the issue

What happens with canard designs (unfortunately) is the fuselage can contribute to directional instability.  Here also the drag of the forward unit when control is applied would be contributory when the nose was pitched.  Try (temporarily) mounting a large tail on the back, essentially making it a tri-plane.  It may be a directional stability issue.  Rutan Vari-eze design may be a good reference.  Where is the engine mounted?

Aerodynamics characteristics of wing and canard look ok. Anyway according to this paper SD7062 suffers from "hysteresis loop at higher Reynolds numbers". Have you done a rough longitudinal stability study? It might be that the CG is not enough forward.

@sophit that is similar to what John is saying.  I agree.


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How are shock waves formed at fighter jet inlets when the shock wave is already formed at the nose cone?

If the main shock wave is formed at the nosecone of a jet aircraft at supersonic speeds, how is the airflow still supersonic at the inlet of the jet intake? Shouldn't the inlets experience subsonic flow since air downstream of shock will be subsonic?

When it is said that "flow is subsonic behind the shockwave", it is only the component of velocity perpendicular to the shockwave which is actually subsonic.

Image source
In a normal shockwave, this makes no difference - the net velocity is the same as its perpendicular component. If the perpendicular component is subsonic (which it always is), the net velocity will inherently be subsonic.
But in an oblique shockwave, the net airflow velocity behind the shock can be supersonic, even though its perpendicular component is subsonic.
In the picture that you shared, you can see that the shockwave is almost entirely oblique (the normal shock at the nosecone is so small that it cannot even be seen in the picture). Anyways, the net velocity of flow behind this shockwave is still supersonic, which means that it will again produce a shockwave when it interacts with the inlets and other parts of the aircraft.

Maybe you could reuse the image from this answer to explain the perpendicular and parallel components

@sophit  ..done!

In the Schlieren photograph you present, you can see that an oblique (slanted) shock is formed wherever there is a change in the cross-section of the aircraft profile, illustrating Aditya Sharma's point that the flow velocity behind an oblique shock is supersonic. Only after traversing a series of oblique shocks does the flow behind the last such shock become subsonic.
This principle of exploiting a series of oblique shocks to progressively decelerate airflow to below mach 1 was used in the design of the J58 inlet structure on the SR-71. In principle, such a series of progressive shocks (which were positioned between the shock cone tip and the first-compressor stage at the face of the J58) can furnish near-perfect ram pressure recovery (lossless compression), thereby performing almost all the compression work required to drive the afterburner of the J-58 during mach 3 cruise.


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In this flying wing concept, how would the aerodynamics be effected by the presence of the intake?

Here the jet intakes are placed to blend in with the flying wing shape. The section where the intakes are located doesn't have the cross section of a conventional airfoil. Will the area along the jet intakes still contribute in efficient lift production?
EDIT: Similar features can also be found in some real life examples
X47 A

X47 B

Aren't you giving this picture a bit much credit? It's not so much a concept as an artist's idea of what "a B-2, but not the B-2" might look like.

@Therac consider it a what if kind of question...

The critical section for lift is the root of the outer wing with its narrow chord. Here a very high lift coefficient is needed which will be limiting for the service ceiling (unless the engines pose this limit). The midsection with its massive chord needs to keep pressure variations from ambient low so the lift distribution does not become distorted.
Intake location here is completely subject to low observable characteristics, and of course there would be better possibilities to place the intakes, but on the upper forward wing they will still work acceptably without spoiling the radar returns too much.

A simplified, maybe overly so, explanation is, that during cruise the intakes will have very little effect on the aerodynamics.
The intakes are usually designed such, that during cruise conditions they simply "scoop" in a slice of air that is the size of the cross section of the intake: There is no suction or overflow at the intake during cruise.
That will, of course be different during takeoff and descent. During takeoff, at slow speed the high rate of in flow will cause low pressure around the intake, and during descent vice versa.
The B-2 has additional scoops behind on the normal inlets to mitigate the low pressure at the intake during takeoff.


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How can I design a slotted flap?
I have not been able to find satisfactory methods on design of slotted flaps, or flaps in general. Most of the answers are about eyeballing or experience. Can you guide me on what could be good starting points.
The images below are what I have design using internet advice, I would like your comments on it.

EDIT: Some information about the design considerations, The airfoil is S2091 with a chord of 40cm and a wingspan of 3.2 metres operating between 0.3 to 1.3 million Reynolds number. Its meant for a UAV that weighs around 35kgs. Without flaps the landing speed is around 18 m/s (65kmph) which I want to lower with use of slotted flaps.
So far I am using thumb rules like having flap roughly the shape of an airfoil and their relative position to trailing edge of airfoils as a given percentage of chord length. But hit and trial doesn't feel like an optimal approach to designing and I seek to understand how others approach this problem. I can design the mechanisms and structures myself, I need help with the aerodynamic design.

Your question is way too broad.  The example looks good to me, voting to close unless you can be more specific about the actual design problem you have encountered.  (I.e. are you asking about aerodynamic shape, structure of the rails, actuation method, etc.?)

I'm asking what is a good starting point in designing a flap, aerodynamically of course. Stop shooting down questions for no good reason that would actually benefit people.

Sorry, if the aerodynamics tag was there before I didn’t notice it.  Still, I think it’s a pretty broad question... weight and speed targets seem like a  logical starting point.

It is a question on aircraft design, which needs to consider all of aerodynamics, structure, actuation etc. Have added the aircraft-design tag and voted to keep open.

I agree with Michael, this is a very broad question. Answerable, yes, but without no info about what kind of a plane is being designed here, the answer will be very general in nature.

Abbot & Von Doenhoff, Theory Of Wing Sections. A classic from the 1950s, containing a chapter on design of high lift devices. Chapter 8 contains aerodynamic design information that may be of interest

Just wanted to ask a quick question on how you designed the section of your drawing between the main body of the aerofoil and the actual slotted flap (section on the right side of the dotted line). Was there a youtube video that you used? Thank you

From the Help Center: "Your questions should be reasonably scoped. If you can imagine an entire book that answers your question, you’re asking too much."  I'd narrow that by saying instead "a full chapter in a textbook" -- and this question is pretty much exactly there.  See the recommendation by @Koyovis above.

see https://aviation.stackexchange.com/questions/99324/how-is-the-geometry-of-slotted-flaps-and-slats-designed


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How is CFD used in aircraft design? Qualitative vs Quantitative?
I have learnt how to do CFD analysis but really I have little idea on what to do with the data to improve/optimise the aircraft design. One approach I can think of is to run lots of analysis for different geometries and compare the results. Another approach may be to visualise flow and get insights from it to proceed to the next iteration.
What I want to ask here is how it is done in the industry. Do major companies just make a lot of small changes and compare thousands of analysis or are there human minds involved after each design iteration using their insights to optimise the aircraft.
edit: How do smaller companies do the same?

This answer might be an interesting read for you. It probably doesn't fully answer your question, but it's a nice example for how Boeing uses CFD.

@Bianfable Thanks. Say one doesn't have access to those codes nor the supercomputers for such calculations. I would like to know how does one proceed then..

Try looking here https://aviation.stackexchange.com/questions/51168/how-to-identify-load-paths/51180#51180

It really depends on what your objectives are. For design optimization, the currently available automated optimization methods (Adjoint and other topology methods) are not as good as a well experienced Aerodynamicist. There has been a lot of effort in improving these tools and creating more complex design optimization tools (such as neural net optimizers), however these tools do not "understand" the fluid flow and simply try and chase whatever metric is being targeted. A good aerodynamicist will be able to make informed decisions based on their interpretation of results in order to iteratively improve the design.
Simulating a large number of different designs and trying to compare the results can be an inefficient use of resources. This method can be useful when comparing a range of options on a similar concept (think simulating 5 options of winglet height to find the optimal). If changing multiple things, it is impossible to work out which change was the most beneficial - the general best practice is to just change one thing at a time.
Also, running simulations across a comprehensive range of roll, pitch, yaw & wing configurations is extremely expensive. A good Aerodynamicist will be able to optimize using simulations at one or a small number of operating conditions and be able to understand the impact of their changes across the entire operating envelope. An automated optimization method would have to simulate many more operating conditions in order to properly evaluate and compare design iterations.
Once a design starts to become more mature, simulating more operating conditions is needed to understand performance across the whole envelope. If the design performs poorly at certain conditions, or if any of the performance characteristics transition too abruptly, the design can then be improved to minimize this unwanted behaviour.
Understanding the fluid flow, and knowing how to optimize the design to meet requirements are the hardest skills to develop. This is what sets good and great Aerodynamicists apart.

In the industry CFD is used for both of the purposes you describe.
CFD simulations are run on existing and proposed designs in order to understand the flow over them, and then based on that understanding to propose design changes or make other alterations using "human minds".
Sometimes the same simulation is run on a variety of different designs, in order to establish which is most efficient.
The two approaches can be intermixed. For example a simulation may be run on a variety of designs to identify the few best ones. Then there may be additional analysis on those best designs, including visualization of the flow, analysis at higher resolutions, analysis of the design with different flow conditions. There may be additional analysis on designs that were expected to be good but weren't, in order to understand why they were good. The design of these analyses is part of what makes a good CFD egineer.

The quick answer is it depends.
As for the small businesses doing aircraft design generally sweeps or trades will be run on a low to mid fidelity multidisciplinary design code to map the design space.  There are many many factors that go into an optimized design, such as what is the purpose/mission for the aircraft, what are design considerations such as it needs to be quiet, carry a certain payload/passenger capacity, fly in specific conditions etc.  Once the larger space has been explored for general configurations, designs identified as more optimal will be more closely analyzed.  This is where CFD generally comes in.  A more refined design will necessitate a higher level of fidelity analysis.  Now you may be running CFD to get more accurate aerodynamics for specific conditions - acoustics, drag reduction, complex geometry etc conditions that just can't be captured at a lower level of accuracy.
This is an example of the aerodynamic design process from the small company I spent 6 years doing aircraft design at.  It's only one example but it gives you an idea:

CFD is an art, grid generation is an art, aircraft design is really a craft.  It is not turn key, there are general rules of thumb and procedures to follow but in my opinion it would be extremely difficult to remove the human factor from this process if you are trying to achieve an optimal design.


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Why aren't grid fins used as wings for low speed aircraft?
Grid fins are used "commonly" in aerospace - usually for things like control of guided bombs or spacecraft. The most obvious example is SpaceX's use on the Falcon 9, but the US and Russian militaries have been using them for decades.
My understanding is that there's a tradeoff with grid fins. They offer greater control authority (and/or lift) in a more compact space, and require less torque to turn/reposition than the equivalent wing would. The downside is that they have a lot more drag, and are almost useless at trans-sonic speeds.
My question is: Why aren't they used as wings (to produce lift) on low-speed aircraft and RC drones? I was unable to find any examples of grid fins being used as wings or lifting surfaces. It seems like if it were possible to create more lift in less space someone would have used it in a design. Am I missing something obvious here? I'm guessing yes. :)

Are you really asking to use grid fins as wings (i.e. to produce lift)? Or did you actually mean to ask about using grid fins as control surfaces?

Yes, I mean using grid fins as the primary/only lift mechanism. I've updated the question accordingly.

Like this https://upload.wikimedia.org/wikipedia/commons/thumb/c/c7/1907_flying_machine.jpg/1280px-1907_flying_machine.jpg ?  From https://en.wikipedia.org/wiki/Horatio_Frederick_Phillips .

Haha yes, like that, although I imagine there's a lot of improvement possible with real airfoils, etc. Are there examples of extreme multiplanes like that in the modern era?

At low speeds, at very small overall width, a heli might outperform any fixed-wing arrangement.

No, not as/on aircraft, and there are good reasons.  They're useful elsewhere, and should make good kites.
Prototype multiplanes flew, quadruplanes reached production, triplanes served operationally in WWI, and biplanes were retired from almost all uses (military and civilian) by early WWII. The biplanes that still fly for purposes other than display are used for low-speed work.
As speeds increase - and we almost always want to get people and goods somewhere quickly - the power lost to drag increases dramatically, in proportion to the cube of the airspeed.  Any construction that increases drag is going to limit speed, payload, range or probably all of them for a given engine power. Biplanes (and to an even greater extent multiplanes) are inefficient at lifting because of the interference between the airflow over the lower wing and under the upper - now squeeze the wings closer together so you can fit more, and this gets worse.
The surfaces (or struts/wires) perpendicular to the lifting surfaces, making up the grid fin or the bracing of a multiplane don't contribute lift but do contribute drag.
Grid fins are sometimes useful: they're compact, more so if the aspect ratio means they can be stowed as by SpaceX and on missiles/glide bombs.  SpaceX use the grid fins in landing, where drag isn't a downside so long as it doesn't cause too much heating.

Thanks, that's super helpful. Just for my own interest, what if we didn't care at all about moving quickly, meaning we could operate as slow as necessary to make the "wings" reasonable? Might things make a bit more sense in that context?

Yes, if the desired speed is arbitrarily low. But if you can't fly into a normal headwind you can't go anywhere, so you'd be looking at a kite rather than an aircraft. And the grid would likely be horizontal lift members with vertical bracing, rather than diagonals. Some of the wind energy kite prototypes approach that, but even they seem to be heading towards well-separated wings


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Did C-124 loadmasters receive parachute and water survival training?
When the position of "loadmaster" was created for the Douglas C-124 Globemaster II, did prospective USAF loadmasters attend the parachute training and water survival school at the 3904th Training Squadron at Fort Carson, Colorado? I believe this school later moved to Stead AFB, Nevada, later the Naval Air Station at Pensacola, Florida (until 2015), and now Fairchild AFB, Washington.

While not a direct answer to the specifically pointed question - the Evolution of the Loadmaster career field: https://www.airforcemag.com/article/1211evolution/

Probably not.  Because if they do now they didn't in 1984, and the globemaster is a lot older than that.


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What do the Boeing 727 airstair tests tell us about D B Cooper?
For the first time in the public domain, the FBI has released one of the technical reports on the Boeing Company’s tests, in March 1964, of the Boeing 727 with the air stairs down in flight. The report in question is numbered and titled D6-7771: “Flight Characteristics with Aft Airstair Down – B”. The airplane used was the second prototype, tail number N72700.
The FBI document is at https://vault.fbi.gov/D-B-Cooper%20/d.-b.-cooper-part-71-of-71/view. The Boeing report is on pages 333-356.
I extracted the following key passages:

Purpose of Conditions Performed: Simulate a failure of the aft airstair up-latch in flight, evaluating the characteristics of the extended airstair and the effect on airplane performance and handling.
The airstair was extended both by allowing it to freefall and by utilizing normal hydraulic power. These extensions were made with the airplane trimmed for level flight at 125 knots and flaps at 25 degrees.
The airstair extended a nominal 8.5 degrees when allowed to free-fall and it required only approximately one-tenth unit of nose-down trim to compensate for the stair.
The stair extended a nominal 13.5 degrees with hydraulic power and this caused approximately a three-tenths unit nose-down trim change.

On November 24, 1971, the hijacker of Flight 305 departed the airplane (a production 727-51, tail number N467US) via the aft airstair, at an indicated air speed of between 160 and 170 knots.
On January 6, 1972, the FBI and the US Air Force flew the same airplane (N467US) in an attempt to replicate the hijacker's jump. On releasing the airstair without hydraulic power at IAS of 150 knots, they found that it rotated through 20 degrees. On placing a person's weight on the bottom step of the airstair, they found that the stair rotated through 35 degrees (vs. fully retracted position). I have not seen any document stating whether they ever used hydraulic power.
Can we conclude that the hijacker of Flight 305, using the control lever without hydraulic power, would have been able to lower the stairs by at least 20 degrees? and if engaging hydraulic power, by significantly more than 20 degrees?
Alternatively are there any known differences between the prototype N72700 and the production 727-51, such that the airstair might be easier to extend on the 727-51? Documents in the public domain indicate that the 727 hydraulic system generally had an operating pressure range of 2800-3100 psi, but I have not found documents specific to the prototypes or the 727-51 model.

could you please provide a link to the document?

I don't understand your question, that report is about airplane handling characteristics, not what it was like to jump off the airstair.

DB Cooper would not have been adjusting the trim, and the plane probably wouldn't have been at 125 knots.

Yes. The result suggests that the hydraulic actuators for the door stall out at that extension, that is, they can't push any harder against air loads, probably by design by limiting extension pressure to the hydraulic cylinders in some way if that was necessary, and the test was to demonstrate that the airplane is manageable in that condition (uncommanded door extension), which is only 5 degrees beyond the free-floating position.
The question becomes, could Cooper crawl out through a 13-14 degree opening while wearing a parachute? Not a very big gap, especially with a military style backpack emergency rig on your back, but apparently, yes.

The tests say nothing about DB Cooper the man, his motives, or whether he survived the jump, but they do provide some insight into the experience of exiting a 727.
However you might envision it happening, he wouldn’t have just walked to the bottom of the stairs and done a swan dive since they cannot be fully extended in flight.  The illustration below from the included link depict the stairs mostly extended, but in reality they would have been closer to an aerodynamic  trail position, or roughly horizontal.  (the accompanying article also has some insight into his instructions for the pilots to slow down and deploy flaps that I was previously unaware of…)

Article picture was taken from
It is difficult to know exactly how much room is available at that extension, but it is likely that he had to crawl along the stairs to squeeze through the exit at the far rear.  The still photos below are taken from the linked YouTube video.  It is difficult to estimate the exact angle, but I attempted to pause at roughly 14 degrees, which as you can see is roughly parallel to the fuselage and would offer much less drag than further extension.

YouTube video of stairs operating
In the second photo taken from the interior you can get at least some perspective of the greatly reduced amount of space there would be in a partial extension.  Snug, and not enough room to standup, but enough to get out.

Finally, 3000 PSI is pretty standard system pressure for aircraft hydraulic system.  It is highly unlikely that an aircraft in normal passenger service would have a system with boosted pressure or a different actuator arrangement intended to allow greater extension than normal.  Unless you can make a positive connection between a surplus airframe modified for jumping that the airline leased and subsequently never modified, and was in service on this particular route on this day, it is "reasonable to conclude that via the control lever, with or without hydraulic power, he would have been unable to lower the stairs by more than 13.5 degrees."

Hmm.  The question, possible answers and comments seem DB Cooper-related.  I don't have a source for this but I have personally met skydivers who rented a 727 and intentionally skydived out of the back without any problems.  I believe it was done more than once by various groups of skydivers in 1980s and 1990s.